1. Field of the Invention
The present invention relates to pressure-fed hybrid rockets. More particularly, the present invention relates to a pressurization system for forcing oxidizer into a hybrid motor of a pressure-fed hybrid rocket, to a novel ignition system for pressure-fed hybrid rockets, and to a novel oxidizer vaporizer system for pressure-fed hybrid rockets.
2. General Background of the Invention
A pressure-fed hybrid rocket requires a pressurization system for the oxidizer tank in order to force the oxidizer out of the oxidizer tank and into the hybrid motor. Classical forward-end injection hybrids typically have a chamber pressure in the range of 300 to 700 psia. Therefore, the oxidizer tank ullage pressure for a pressure-fed hybrid rocket is in the range of 400 to 800 psia. Pressurization system size and weight impacts vehicle size and weight, therefore affecting the payload attainable.
A reliable, safe ignition system is desirable for pressure-fed hybrid rockets. Prior to the present invention, there has been no such system.
There are various types of pressurization systems, the closest prior practice to the system of the present invention being an inert gas plus combustion pressurization system. The heat source for an inert gas plus combustion pressurization system in the past has been in the form of either a catalyst or gas generator. These types of heat source tend to be relatively expensive and inefficient in heating the pressurant, whereas the hybrid helium heater concept of the present invention provides a low cost pressurization system.
Bradford et al., U.S. Pat. No. 5,119,627 contains a good discussion of pressure-fed hybrid rockets, and is hereby incorporated by reference. In the paragraph spanning columns 1 and 2, the '627 patent describes a prior art pressurization system for pressure-fed hybrid rockets which includes high-pressure non-flammable gas fed into the oxidizer tank through a regulator or throttle valve.
Bradford et al. describes a system including a tank containing helium gas under pressure, a tank containing a liquid oxidizer, for example, and a tank containing solid fuel. The helium gas under pressure travels through a conduit into the liquid oxygen tank and expels liquid oxidizer into the solid fuel tank. In Bradford et al., however, the gas is stored at ambient temperature and is not heated prior to being introduced into the oxidizer tank.
Schuler et al., U.S. Pat. No. 5,099,645, discloses a liquid-solid propulsion system including a solid fuel tank and a liquid oxygen tank wherein the liquid oxygen is heated up to become gaseous before it enters the tank containing the solid fuel, and some of the gaseous oxygen is purportedly used to push more liquid oxygen out of the liquid oxygen tank.
Schubert et al., U.S. Pat. No. 3,595,020, discloses a rocket which uses solid fuel. However, the fuel and the oxidizer are both solid. The gas produced by the combustion of the solid fuel and oxidizer pushes on a piston which causes a liquid oxidizer to be ejected from a chamber into a combustion chamber where burnable gases produced by the combustion of the solid fuel and solid oxidizer are burned.
Knuth et al., U.S. Pat. No. 5,101,730, discloses a gas-fed hybrid propulsion system which uses a turbopump to pump oxygen into contact with a solid fuel rocket.
An adequate pressurization system is just one of the requirements of a successful pressure-fed rocket. A reliable ignition system is also important, as is smooth, predictable, reliable combustion.
The fundamental principle that allows a hybrid rocket to burn is that in steady state operation, the fuel surface is constantly generating a melt layer which in turn generates vapor as more heat is added or the heat causes the fuel to sublime directly to vapor from solid phase. In the early days of hybrids, room temperature oxidizers with low heat of vaporization were used in experiments with rubber and plastic to develop the hybrid combustion theory. The considerations when using a cryogenic oxidizer like LO2 (liquid oxygen) are substantially different. The extreme cold of the LO2 can freeze the surface of the fuel, totally eliminating its ability to generate fuel. This causes low frequency oscillations in chamber pressure because of the time required to transfer heat to the cooled surface when the combustion zone travels back over a "quenched" fuel zone. That finite amount of time limits the maximum speed that the fuel vapor can be re-generated after the fuel has been quenched. Higher frequency oscillations are found when using a room temperature liquid oxidizer with a low heat of vaporization or one in gaseous form.
Many of the prior art hybrid rockets use hypergolic liquids to start them. Campbell (U.S. Pat. No. 3,116,599) uses a liquid for starting multiple times.
The hypergolic liquids used by prior art hybrid rockets are very hazardous. They spontaneously burn when exposed to oxidizer (or else they would not work) and most of them spontaneously bum when exposed to air (the oxygen in air), resulting in substantial safety risks and concerns. Any leakage of the liquid fuel will ignite the launch vehicle. Special storage and handling techniques are required to handle the liquids, resulting in substantial costs.
Hybrid propulsion rockets are safe and non-explosive (because the fuel is in solid phase and cannot explode and detonate). To employ the ignition method of the prior art, a quantity of hazardous liquid must be added to the launch vehicle, negating the principal advantage.
The hypergolic fluids can only burn if they are vaporized immediately and mix with vaporized oxidizer. The real problem is when using liquid oxygen. The cryogenic temperature of the LO2 naturally suppresses the reaction rate (which roughly doubles every 10.degree. C.). Further aggravating the problem, hypergolic liquids typically have a relatively high freezing temperature. Triethyl aluminum, for instance, freezes at 59.degree. F. The introduction of a large quantity of LO2 into a hybrid creates a region in which the temperature is effectively cryogenic. Initially, the LO2 comes in contact with ambient (i.e. warm) surfaces and vaporizes. Once the surface heat has been removed, the spray has little opportunity to vaporize. In such a cold predominately liquid LO2 stream, the hypergolic liquid cannot be vaporized (as it requires thermal energy to change phase), in fact, it is more likely that it is frozen (since these liquids have a high freezing point). The residence time in the head end is very small, and the frozen hypergolic fuel does not have enough time to vaporize, mix and burn. The frozen hypergolic fuel is then swept down into the port of the hybrid motor where it eventually thaws, vaporizes and burns with oxygen.
As the cold oxygen travels down the port, it will eventually be vaporized and will allow continuous combustion with the hybrid grain. The more LO2 that is fed down a given unit area of port (oxygen flux), the farther down the port the steady state combustion zone will be. The farther down the port the steady state combustion zones, the higher the potential amplitude of unstable combustion response. If the distance is zero, the unstable amplitude due to this mechanism is also zero.
Analytical methods show that, even for low oxygen flux, the zone can be located deep into the port. Combustion instability is caused by the oscillation of the combustion zone forward, which generates more fuel to burn (surface area exposed to combustion), resulting in a higher pressure, and rearward which allows the cold oxygen to quench the fuel surface and eliminate its ability to generate flammable fuel vapor. The highest possible frequency of oscillation due to this phenomena is limited by how quickly the fuel surface can be re-heated as the combustion zone returns forward. The inherent stability of the hybrid combustion process (i.e. no destructive high-frequency oscillations that are catastrophic in liquid motors) is due to this time delay.
Various patents of interest relating to ignition systems and to vaporizing liquid oxidizer are discussed below.
Bradford et al., U.S. Pat. No. 3,518,828, discloses a hybrid rocket ignition system which directs the fuel-rich combustion products of a small solid propellant igniter into the recirculation region 32 around an oxidizer spray 33 in the rocket case. Once the solid propellant 43 ignites, it cannot be extinguished until it is consumed. This negates one of the most important safety features of hybrid rockets--the ability to start and stop the rocket's combustion at will. Also, solid propellant is relatively dangerous and can be ignited accidentally.
Holzman, U.S. Pat. No. 3,295,323, discloses a means for vaporizing LO2 inside a rocket case; however, the LO2 is vaporized while in tubing.
Avery, U.S. Pat. No. 3,136,119, discloses an ignition system for a solid-oxidizer hybrid rocket which directs air or oxygen into or around an incoming liquid fuel stream (see especially col. 5, lines 22-25)--it is not in direct contact with the combustion gases.
Muzzy, U.S. Pat. No. 3,782,112, discloses a system for gasifying and aerating a liquid oxidizer before it enters a hybrid rocket. While the details are sketchy, it appears that solid propellant 18 somehow generates a "very hot gas" which passes through openings 21 and 22 and mixes with the liquid oxidizer. However, note that this mixing of the "very hot gas" and the liquid oxidizer occurs prior to entry of the oxidizer in the combustion chamber. Also, as does the apparatus of Bradford et al., it uses a solid propellant, which is relatively dangerous and can be ignited accidentally. Further, the solid propellant cannot be shut off and restarted.
Greiner, U.S. Pat. No. 2,996,880, discloses a rocket which is ignited using electrical resistance (either the solid fuel or a separate resistor can be used to produce the necessary resistance).
Mangum, U.S. Pat. No. 3,340,691, discloses a system in which a solid propellant is ignited electrically, then both pressurizes liquid and mixes with the liquid and directs the liquid against a solid barrier with which the liquid reacts.
Lai et al., U.S. Pat. No. 3,613,583, discloses a flare which uses methane as an ignition fuel, which would be dangerous in a rocket.
Bruun et al., U.S. Pat. No. 3,613,583, discloses a monopropellant thrust or which includes tubes in a thermal bed. Hot gases pass through the tubes to help heat the bed to enhance decomposition.
Schuler et al., U.S. Pat. No. 5,099,645 (also discussed above), discloses a liquid-solid propulsion system including a solid fuel tank and a liquid oxygen tank wherein the liquid oxygen is heated up to become gaseous before it enters the tank containing the solid fuel. Further, in Schuler et al., the liquid oxygen is heated in a heat exchanger, and does not come in direct contact with the exhaust stream of the hybrid heater.
Knuth et al., U.S. Pat. No. 5,101,730 (also discussed above), discloses a gas-fed hybrid propulsion system which uses a turbopump to pump oxygen into contact with a solid fuel rocket. In Knuth et al., the LO2 and a fluid fuel are mixed in a preburner 7 to heat and gasify the oxygen.